This invention relates generally to gas turbine engines and, more particularly, to methods and apparatuses for reducing turbine rotor temperatures.
Gas turbine engines typically include a compressor, a combustor, and a high-pressure turbine. In operation, air flows through the compressor and the compressed air is delivered to the combustor wherein the compressed air is mixed with fuel and ignited. The heated airflow is then channeled through the high-pressure turbine to facilitate driving the compressor. Moreover, during operation, un-cooled high-pressure turbine blades may transfer heat from the turbine blades, at gas path temperature, through the shank, and by conduction and/or convection, to the high-pressure turbine disk. Furthermore, cooling flow lost due to shank leaks my allow combustion gases to enter the cooling circuit, exposing the turbine disk to combustion gas temperatures. As a result, the turbine disk is exposed to high temperatures which may thermally fatigue the turbine disk.
To facilitate preventing damage that may result from turbine disk exposure to high temperatures and possibly combustion gases, at least one known gas turbine engine includes an internal cooling circuit to facilitate cooling the turbine disk. More specifically, cooling air is channeled along a forward face of the disk from a radially inner portion of the disk along a substantially linear path to a radially outer portion of the disk. However, channeling the cooling air linearly along the face of the rotor disk may not effectively cool the disk. Moreover, various fasteners and/or blade retainer pins within the cooling flowpath create undesired temperature rise due to windage, which may further reduce the ability for the cooling air to effectively cool the turbine disk.